Durable refractory ceramic coating

ABSTRACT

An article used for thermal protection includes a base structure having at least one surface. The base structure is made of a ceramic oxide material. A silicide coating is formed on the at least one surface of the base structure. The coating comprises a refractory metal and silicon, which together form a silicide. The coating is at least partially diffused into the base structure at the at least one surface.

BACKGROUND OF THE INVENTION

[0001] The present invention relates generally to refractory materials,and more specifically, to insulating materials having refractory ceramiccoatings. A base structure is coated with a refractory silicide coating.The coating is made of a refractory metal, i.e., those having a meltingpoint above about 1,650° C., and silicon. The combination of thesematerials forms a “silicide” coating.

DESCRIPTION OF THE RELATED ART

[0002] High temperature environments such as those found in atmosphericreentry, jet turbine combustion, or rocket propulsion, necessitate theuse of thermal protection systems that provide oxidation protection,high emissivity, and resistance to mechanical damage. One example ofsuch a system is the tile used to cover the outer surfaces of theunderbody and wings of NASA's Space Shuttle Orbiter. Each tile is alightweight, fibrous, silica-based rigid fibrous insulation unit with anominal density of nine (9) pounds per cubic foot (pcf). The tile ismade by the assignee of the present invention and designated as “LI-900.

[0003] A variation of LI-900, called LI-2200, and likewise manufacturedby the assignee herein, is a twenty-two (22) pcf version of the LI-900that offers improved strength at a sacrifice in weight.

[0004] To improve strength, a new class of rigid reusable insulation wasdeveloped. These consist of the following composite ceramic materials:FRCI, AETB, and HTP. A further variation of LI-900, made by NASA's AmesResearch Center and designated “FRCI-12,” consists of a blend of silicaand aluminoborosilicate fibers. FRCI-12 has a density of twelve (12)pcf. All three of these materials are currently qualified for use on theShuttle Orbiter Fleet.

[0005] “HTP,” which stands for “high thermal performance,” refers to anew class of lightweight ceramic material introduced by the assigneeherein around 1982. Basically, this high-strength insulation is producedby fusing silica and alumina fibers together. The insulation is producedto a number of standard densities at a standard composition oftwenty-two (22) percent. The HTP family of insulants has yieldedimprovements in strength and maximum temperature capability relative toearlier generations of ceramic insulation. Also, at about the time HTPwas introduced, NASA (Ames Research Center) introduced the rigid fibrousinsulation material known as “AETB.”

[0006] Coatings have been used in conjunction with refractory metals,such as tantalum, niobium, and molybdenum, to protect the underlyingmetallic structures from oxidation and erosion experienced in hightemperature propulsion environments. Silicide coatings have been used inthe past for such purposes.

[0007] The TPS tiles noted above have in the past been protected byapplication of reaction cured glass (RCG) coatings. These coatings weredeveloped in the early 1970's for the LI-900 class of thermal insulants.RCG is composed mostly of silica with a small amount of siliconhexa/tetraboride added as a blackening agent and a fluxing agent. Thecoating is applied as a 8-12 mils thick layer onto the surface of aceramic tile. As a surface coating, RCG has relatively poor resistanceto impact; as a silica-based system, RCG's maximum temperaturecapability is limited to its softening point of about 2,700-2,800° F.

[0008] In 1989, NASA's Ames Research Center developed an insulationproduct called toughened uni-piece fibrous insulation (TUFI). Thecoating was still silica-based, but it contained about twenty (20)percent molybdenum disilicide as a blackening agent. TUFI productsrepresented an advancement in the state of the art because the coatingis applied as a surface impregnation, meaning that it became commingledwith the fibers of the insulation tile near the surface region. Theresultant fused coating is a fiber reinforced glass which is much moredurable than the RCG coating. As a silica-based coating, however, theTUFI product has the same upper temperature limit as RCG, i.e.,2,700-2,800° F. TUFI has been successfully applied to FRCI, HTP, andAETB.

[0009] Refractory metal coatings have been used in ceramic applications.For example, U.S. Pat. No. 5,413,851 to Storer describes a ceramiccarbon fiber coated with a refractory metal or metal-based ceramicmaterial. The refractory metal materials used include molybdenum,tantalum, tungsten, niobium, oxides of aluminum, yttrium, zirconium,hafnium, gadolinium, titanium, erbium and other rare earth metals. Thefibrous materials that are coated include alumina, alumina-silica, andalumina-boria- silica. The coatings are used to enhance strength.

[0010] U.S. Pat. No. 4,530,884 to Erickson et al. describes aceramic-metal laminate which is used as insulation in high temperatureenvironments. The composites described therein have an inner ceramiclayer and an outer metal layer and an intermediate interface layer of alow modulus metallic low density structure. These composites areprincipally used as turbine blades of gas turbine engines.

[0011] U.S. Pat. No. 5,863,846 to Rorabaugh et al. describes a ceramicinsulation used in aerospace applications in which a slurry is moldedfrom ceramic fiber to form a soft felt mat which is impregnated with asol prior to drying. The mat is exposed to a catalyst that diffuses intothe mat and causes the sol to gel.

[0012] U.S. Pat. No. 5,814,397 to Cagliostro et al. describes ceramicmaterials used in space re-entry vehicles in which silica coatings areformed on fibrous insulations. U.S. Pat. No. 5,079,082 to Leiser et al.describes a porous body of fibrous, low density silica-based insulationmaterial that is impregnated with a reactive boron oxide-containingborosilicate glass frit, a silicon tetraboride fluxing agent, and amolybdenum silicide emittance agent.

[0013] A continuing need exists for improved lightweight thermalinsulation materials that are temperature resistant and physicallydurable. In particular, improved coatings, such as those describedbelow, are needed for all of the advanced rigid, fibrous insulationmaterials described above, such as FRCI, HTP, and AETB.

SUMMARY OF THE INVENTION

[0014] An object of the present invention is to provide an insulativematerial that exhibits oxidation protection, high emissivity, andresistance to mechanical damage.

[0015] Another object of the present invention is to provide aninsulative material that is capable of withstanding high-temperatureenvironment, including those associated with atmospheric reentry, jetturbine combustion, and rocket propulsion.

[0016] Still another object of the invention is to provide a silicidecoating that exhibits higher temperature capability than silica-basedcoatings.

[0017] These and other objects of the invention are met by providing athermal protection system comprising a base structure having at leastone surface, the base structure being made of a ceramic oxide material,and a coating formed on the at least one surface, the coating comprisinga refractory metal and silicon and being at least partially infiltratedinto the base structure at the at least one surface.

[0018] The foregoing features and advantages of the present inventionwill be further understood upon consideration of the following detaileddescription of the invention taken in conjunction with the accompanyingdrawings.

BRIEF DESCRIPTION OF THE DRAWINGS

[0019]FIG. 1 is a perspective view of a thermal insulation systemaccording to the present invention;

[0020]FIG. 2 is an enlarged side elevational view of the thermalinsulation system of FIG. 1;

[0021]FIG. 3 is an photo-micrograph, magnified at 50×, of the thermalinsulation system of FIG. 1;

[0022]FIG. 4 is a graph showing impact energy vs. failure rate of thethermal insulation system of FIG. 1, compared to that of a prior artinsulation material; and

[0023]FIG. 5 is a photograph showing two samples of material afterexposure to high temperatures, in which a sample of the presentinvention exhibited high temperature resistivity.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0024] Referring to FIG. 1, a thermal insulation system is generallyreferred to by the numeral 10. The system can be an insulatingstructure, such as a tile or any other object shaped and sized toperform a specific function. The polyhedron shape of the system 10 waschosen for illustration purposes. Virtually any shape is contemplated tobe within the scope of the present invention. Of course, a typicalsystem may employ numerous, similarly shaped objects, such as in thetiles used on the outer surfaces of NASA's Shuttle Orbiter. For example,the thermal protection system 10 could be a thermal insulation tilehaving the dimensions of six (6) inches by six (6) inches by two (2)inches. The tiles may be flat or profiled to adopt any desirable shape.

[0025] The thermal protection system 10 includes a base structure 12having at least one outer surface. The base structure 12 is preferablymade of a ceramic oxide material. Preferred ceramic oxide materialsinclude those advanced examples discussed above. While other materialscan be used, the base structure is preferably a ceramic oxide, forexample: zirconia, alumina, and/or silica/alumina blends.

[0026] A coating 14 is formed on one or more of the surfaces of the basestructure 12. The coating 14 is made of a refractory metal and siliconand is at least partially diffused into the base structure 12 at the atleast one surface. The refractor silicide coating employs the samesurface impregnation strategy as the silica-based TUFI materials used inthe past, but substitutes a composition that is nearly 100% refractorymetal silicide. The coating 14 is formed from any of the refractorymetals, meaning those elements with melting points in excess of 1,650°C., in combination with silicon. This combination produces a “silicide”coating. Examples include molybdenum silicide, tantalum silicide andniobium silicide.

[0027] A particularly preferred material for the coating 14 ismolybdenum silicide. Molybdenum disilicide has a melting point that isapproximately 350° F. higher than silica. The result is that thesilicide coating has higher temperature capability than the silica-basedcoatings used in RCG and TUFI materials. Moreover, since the silicidecoatings described herein are applied as a surface impregnation, thepresent materials are as durable as the TUFI materials. RCG and TUFIsilica-based coatings have a melting point of about 3,180° F. and asoftening point of between about 2,700 and 2,800° F. In contrast, thepresent silicide-based coatings have a melting point of about 3,540° F.,and a softening point between about 2,900 and 3,000° F.

[0028] Any of the refractory metals can be used as a constituent of thecoating 14, including molybdenum, tantalum, niobium, vanadium, chromium,and tungsten, or any other materials that are comparable chemically andwith respect to melting point.

[0029] The molybdenum silicide coating 14 is prepared from high puritymolybdenum disilicide and molybdenum diboride powders. Boron nitride andother forms of boron may be used as an alternative source of boron. Thesilicide provides the desired high melting point and high emissivitycharacteristics. The boride serves as a flux to allow low temperaturefusion of the coating and also plays a role in providing oxidationresistance to the coating. As with RCG and TUFI coatings, these powdersare formulated into a slurry and applied at a pre-specified coatingweight to the surface of the ceramic oxide base structure 12 by eitherbrushing or spraying. The coating is fused in an air furnace at atemperature in the range of about 1,800° F. to about 2,600° F. A morepreferred range is between 2,100° F. and 2,400° F.

[0030] The slurry which contains the constituent powdered materialsincludes a liquid carrier, which is a polymeric stock solution known andused in the art. The heating time in the air furnace can be anytimesufficient to allow fusion and attachment of the coating to the basestructure. A typical time range is between one (1) and two (2) hours.The temperature is likewise selected to effect fusion of the coating,and should be kept below the sintering temperature of the substrate, orbase structure, which for the aforementioned preferred materials isabout 2,600° F.

[0031] The coating 14 can be varied in thickness. “Thickness” refers tothe coating material on the surface and that which has diffused into thebase material, forming an infiltration zone beneath the surface. In aparticularly preferred embodiment, for tiles used on the X-33 spacecraftcurrently under development, the coating 14 has a weight of about 1.0gram per square inch, which provides a coating of about 0.030 inches. Inthis particular structure, the thickness of 0.030 inches represents onlythe surface coating and not the infiltration zone. In general, theinfiltration zone varies depending on the substrate material. Agenerally preferred range of weights is 0.25 to 2.0 grams per squareinch for the surface coating (excluding the infiltration zone). Thecoating durability tends to increase with higher coating weights. Itwill be recognized by those skilled in the art that a broader range ofweights and thickness are possible, depending on the specifications ofthe end product.

[0032] The coating 14 is applied to achieve a true surface impregnationresulting in a silicide coating that is reinforced by the fiber networkof the underlying silica-based rigid fibrous insulation material whichforms the base structure 12. this can be seen in FIG. 3, which is aphotomicrograph of a base structure coated with molybdenum disilicide.The molybdenum disilcide is shown as the lighter colored regions, andthe ceramic oxide material which forms the base structure is shown asthe darker regions. It is evident that the molybdenum disilicide hasdiffused into the base structure. The infiltration zone, referred toabove, corresponds to the distance between the outer surface of thesurface coating and the lower-most area of penetration of the coatingmaterial into the base material. Referring to FIG. 3, the thickness ofthe surface is approximately shown by the letter “A” and theinfiltration zone, which includes the surface depth, is shown by theletter “B.” The result is a highly durable coating.

[0033] Durability of products made according to the present invention isdemonstrated with reference to FIG. 4, which shows impact test resultsfor LI-900 prior art coated ceramic thermal protection tiles, and HTP-6which incorporates the refractory silicide coatings of the presentinvention. The test was conducted as follows: A steel ball having avolume of one (1) cubic centimeter was dropped from a fixed height andimpacted the surface with an energy measured in Joules (along the x-axisof FIG. 4). For each impact energy, five drop tests were conducted andthe fraction of coating failures was noted. The drop height for 0.002Jwas 2.5 cm; for 0.007 the drop height was 8.0 cm; for 0.026J the dropheight was 32.0 cm; and for 0.042 the drop height was 51.0 cm. Coatingfailure was defined as follows: “damage threshold” for RCG coating isthe appearance of a surface crack; “Brinnell failure” for RCG coating islarge area cracking of the RCGT accompanied by crushing of underlyingLI-900 material; and “damage threshold” for LMMS refractory silicidecoating is surface spall. For the latter, the coating does not fail bycracking as does RCG. As is evident from the plotted points, the presentinvention achieved a factor of 25 improvement in impact energy.

EXAMPLE

[0034] To make tiles used for the X-33 spacecraft, a slurry was createdby using 43% by weight molybdenum silicide powder (100 mesh), 7% byweight boron nitride powder (325 mesh), 2.5% by weight liquid stocksolution A, which is a polymer-based binder, and 47.5% denaturedethanol. These materials are combined and jar milled to produce a sprayslurry which has a density of 11.5+/−0.5 pounds per gallon, and aparticle size distribution as follows: Mass % finer Microns 90 5.25 to13.76 50 3.30 to 5.57  10 .51 to 1.50

[0035] The slurry is then sprayed to desired thickness and then dried attemperatures and times specified above.

[0036] Products made according to the present invention were testedagainst the prior art and were shown to have marked increases in thermalperformance and durability. In a comparative burner rig test, refractorysilicide coatings of the present invention were tested against a NASAAmes TUFI+RCG. Both the present coatings and the NASA coatings were madeon a base structure made of HTP 8-22. The TUFI+RCG coating had a weightof 1.07 g/in² and the present coating, a molybdenum silicide(MoSi₂+MoB₂), had a total coating weight of 0.57 g/in².

[0037] Test samples were exposed to an array of torches fueled byacetylene, hydrogen, and oxygen. The temperature was measured by opticalpyrometer and thermocouple. FIG. 5 shows two test samples, one being thepresent invention and the other the prior art, after 6 minutes of heatat 1590° C. The sample on the right shows heavy pitting, while thepresent invention, the sample on the left, survived the high temperatureheating. Other tests have shown that the present coatings can withstandtemperatures in a flame test of up to about 2,950° F.

[0038] While advantageous embodiments have been chosen to illustrate theinvention, it will be understood by those skilled in the art thatvarious changes and modifications can be made therein without departingfrom the scope of the invention as defined in the appended claims.

What is claimed is:
 1. A thermal protection system comprising: a basestructure having at least one surface, the base structure being made ofa ceramic oxide material; and a silicide coating formed on the at leastone surface, the coating being at least partially diffused into the basestructure at the at least one surface.
 2. A thermal protection systemaccording to claim 1 , wherein the silicide coating is molybdenumsilicide.
 3. A thermal protection system according to claim 1 , whereinthe silicide coating is tantalum silicide.
 4. A thermal protectionsystem according to claim 1 , wherein the silicide coating is niobiumsilicide.
 5. A thermal protection system according to claim 1 , whereinthe refractory metal is selected from the group consisting of tungsten,molybdenum, tantalum, niobium, vanadium and chromium.
 6. A thermalprotection system according to claim 1 , wherein the coating includesboron in an amount sufficient to lower the melting point of the coatingand thereby facilitate fusion of the coating.
 7. A thermal protectionsystem according to claim 1 , wherein the silicide coating includesmolybdenum silicide and molybdenum boride.
 8. A thermal protectionsystem according to claim 1 , wherein the base structure is asilica-based rigid fibrous insulation material of predetermined size andshape.
 9. A thermal protection system according to claim 1 , wherein thesilicide coating has a weight of between about 0.25 and 2.0 grams persquare inch.
 10. A thermal protection system according to claim 1 ,wherein the silicide coating is temperature resistant up to about 2,950°F.
 11. A thermal protection system according to claim 1 , wherein thesilicide coating has an impact energy for 50% failure of about 0.02Joules.
 12. A method of forming a coated thermal insulation article,comprising the steps of: forming a base structure from a ceramic oxidematerial; forming a slurry from a liquid carrier constituent and apowdered constituent which contains silicon and a refractory metal;applying the slurry in a predetermined coating weight to a surface ofthe base structure; and heating the coated base structure for a time andtemperature sufficient to fuse the coating to the base structure.
 13. Amethod according to claim 12 , wherein the step of forming a slurryfurther includes adding a flux agent to the slurry to allowlow-temperature fusion of the coating and to further enhance theoxidation resistance of the coating.
 14. A method according to claim 13, wherein the flux agent is selected from the group consisting ofmolybdenum diboride and boron nitride.
 15. A method according to claim12 , wherein the applying step includes applying the slurry by a processselected from the group consisting of spraying and brushing.
 16. Amethod according to claim 12 , wherein the heating step comprisesheating the coated base structure for 1-2 hours at a temperature ofbetween 1,800 and 2,600° F.
 17. A method according to claim 13 , whereinthe powdered constituent is selected from the group consisting ofmolybdenum disilicide, tantalum disilicide, niobium disilicide,molybdenum diboride, tantalum diboride, niobium diboride, and boronnitride, and combinations thereof.
 18. A product made by the process ofclaim 12 .
 19. An article used for thermal protection, comprising: abase structure having at least one surface, the base structure beingmade of a ceramic oxide material; and a silicide coating formed on theat least one surface, the coating comprising a refractory metal andsilicon, which together form a silicide, the coating being at leastpartially diffused into the base structure at the at least one surface.20. An article according to claim 19 , wherein the silicide coating isselected from the group consisting of molybdenum silicide, tantalumsilicide, and niobium silicide.
 21. An article according to claim 19 ,further comprising a diffusion zone disposed between the silicidecoating and the base structure, and being composed of coating materialthat diffuses into the base structure, wherein the silicide coatingincludes boron in an amount sufficient to lower the melting point of thecoating and thereby facilitate fusion of the coating the base structure.